Skip Nav Destination
Close Modal
Update search
Filter
- Title
- Author
- Author Affiliations
- Full Text
- Abstract
- Keyword
- DOI
- ISBN
- ISBN-10
- ISSN
- EISSN
- Issue
- Volume
- References
- Conference Volume
- Paper No
Filter
- Title
- Author
- Author Affiliations
- Full Text
- Abstract
- Keyword
- DOI
- ISBN
- ISBN-10
- ISSN
- EISSN
- Issue
- Volume
- References
- Conference Volume
- Paper No
Filter
- Title
- Author
- Author Affiliations
- Full Text
- Abstract
- Keyword
- DOI
- ISBN
- ISBN-10
- ISSN
- EISSN
- Issue
- Volume
- References
- Conference Volume
- Paper No
Filter
- Title
- Author
- Author Affiliations
- Full Text
- Abstract
- Keyword
- DOI
- ISBN
- ISBN-10
- ISSN
- EISSN
- Issue
- Volume
- References
- Conference Volume
- Paper No
Filter
- Title
- Author
- Author Affiliations
- Full Text
- Abstract
- Keyword
- DOI
- ISBN
- ISBN-10
- ISSN
- EISSN
- Issue
- Volume
- References
- Conference Volume
- Paper No
Filter
- Title
- Author
- Author Affiliations
- Full Text
- Abstract
- Keyword
- DOI
- ISBN
- ISBN-10
- ISSN
- EISSN
- Issue
- Volume
- References
- Conference Volume
- Paper No
NARROW
Date
Availability
1-20 of 619
Test facilities
Close
Follow your search
Access your saved searches in your account
Would you like to receive an alert when new items match your search?
Sort by
Proceedings Papers
Proc. ASME. GT2020, Volume 2A: Turbomachinery, V02AT32A017, September 21–25, 2020
Paper No: GT2020-14378
Abstract
The goal of the paper is to propose modifications to tested flat cascades that will suppress partial distortions of acquired interferograms in the region of blade leading edges. This improvement will allow more accurate determination of actual flow incidence angles of tested blade cascades, in particular for transonic or supersonic inlet flows. Application of physical probes for such tasks is always in question in transonic or supersonic flows. The paper is composed of three main sections: (a) introduction of the test facility, (b) presentation of the problem with examples, and (c) description of the experimental work. Recommendations for future flat cascade investigations is presented in the paper. The first section is devoted to the introduction and description of the High-Speed Laboratory of the Institute of Thermomechanis of the Czech Academy of Sciences. Attention is paid to the unique large-scale interferometer which is one of the principal research instruments here and which is routinely used for investigations of transonic compressor and turbine cascades. The instrument capability is illustrated by a series of images showing evolution of a sonic line in a transonic cascade as a function of the increasing inlet Mach number. The reasoning for the proposed work is presented in the middle section. The major impetus for the work was to understand the observed discrepancies between schlieren and interferometer images while testing highly-loaded transonic compressor cascades. In particular, the main concern is the relatively wide region of increasing pressure in the shock vicinity recorded on interferograms versus sharp shock wave image visible on schlieren images. It was suggested that these discrepancies are caused by deformation of the shock-wave surface by the growth of secondary flow due to the tunnel endwall effects. It should be stressed here that the intention was not to investigate the pattern or the nature of the secondary flows rather. An idea behind this approach is to move the secondary flows out of the region of interferometer imaging. Finally, in the last section the results of the experiments carried out during the course of this work are presented. The experiments were designed to improve understanding of the origins of interferogram distortions. Further intention was to eliminate or at least lessen the level of interferogram distortions due to the combined effects of the boundary layer interaction and the corner-vortex flow. Wedges of a constant vertex angle of 15 deg of various plane shapes were inserted subsequently in supersonic flow of (Mach number 2) and interferograms of the resulting flow pattern were acquired. It was observed that decreasing the wedge span led to clearing the interferograms of the superimposed distortions. This confirmed the decisive role of the end wall effects on the quality of acquired results. The undistorted interferograms of the inlet flow in the region of the shock structure are needed to determine the actual angle of attack of the incoming flow onto the tested transonic cascade. Based on the presented results it is suggested for the future testing of flat cascades to modify the front part of the blades by appropriate side cut-offs to eliminate interferogram distortions.
Proceedings Papers
Proc. ASME. GT2020, Volume 2A: Turbomachinery, V02AT32A013, September 21–25, 2020
Paper No: GT2020-14322
Abstract
Rotating stall is an unsteady flow phenomenon in flow rate reduction process of axial compressor. Single or multiple stall cells can be found in blade passage, rotating circumferentially at frequency ranging from 20% to 80% of compressor rotating frequency. Rotating stall may further lead to surge, then mass flow and pressure ratio of the compressor decrease rapidly while vibration amplitude of bearing fulcrum increases significantly, which would affect safety operation of the aero-engine. This paper reports the study on a multistage compressor with special-shaped aggressive flow path, which is significantly different from the conventional compressor with its meridian flow path midline parallel or nearly parallel to the rotating shaft axis. To study rotating stall and surge characteristics of multistage compressor with inlet distortion, this paper takes a 4.5-stage axial compressor as research object which has special S-shaped meridian flow path. On SJTU compressor test facility, the author carries out multidisciplinary tests on compressor dynamic performance (stator LE/TE and inter-stage fluctuating pressure measured by fluctuating pressure transducers, vibration measured by acceleration transducers, and dynamic stress measured by pasting strain transducers on the surface of stator blade). And in subsequent surge experiment, stall and surge signals at different rotating speeds are successfully captured. With above test data, axial, radial and circumferential instability development characteristics are studied by time domain correlation analysis, frequency domain power spectrum and coherence analysis, which preliminarily reveals the generation, development, propagation and elimination characteristics of rotating stall and surge. Meanwhile, in view of time-frequency characteristics of two instability transition processes, time-frequency analysis method based on Choi-Williams distribution model is adopted to capture instantaneous frequency characteristics. The results show that instability phenomenon exhibits quasi-exponential change in instability transition processes. Compared with traditional methods, time-frequency method is more effective in characterizing strong time-varying instability signals such as rotating stall and surge.
Proceedings Papers
Proc. ASME. GT2020, Volume 2A: Turbomachinery, V02AT32A067, September 21–25, 2020
Paper No: GT2020-16031
Abstract
In this paper, steady and unsteady CFD have been used to investigate stall inception for a modern low pressure ratio transonic fan. The computational results are validated against measurement data from a high-speed test facility. CFD validation was approached as a blind test case. The results shows good agreement between the experiments and computations. Stall is triggered by growth of a suction surface separation behind the shock around the mid-span of the rotor blade. As the fan is throttled the separation grows leading to increased blockage in the blade passages. At the point of instability the separation grows further, locally increasing incidence and leading to the formation of a stall cell. It is shown that changes to the tip leakage flow leave the stall inception mechanism unaffected. A computational case with a suction surface slip patch between 25–75% span shows that the reduction in blockage around the mid-span increases the stall margin by 25%. This demonstrates that for cases with mid-span initiated stall it is important to consider the flow away from the tip as well as the flow in the tip region. A redesigned fan is used to illustrate that design changes around the mid-span can be effective to improve flow range. The redesigned fan increases stall margin by 6.7% while maintaining the design point efficiency within 0.1%.
Proceedings Papers
Proc. ASME. GT2020, Volume 2D: Turbomachinery, V02DT36A013, September 21–25, 2020
Paper No: GT2020-14897
Abstract
The presented work investigates the tone noise of three booster stages designed for a high bypass ratio turbofan. Calculations were performed at the approach operational conditions. The frequency domain numerical method of multistage turbomachines tone noise simulation, developed in CIAM (Central Institute of Aviation Motors) and implemented in 3DAS (3 Dimensional Acoustics Solver) in-house solver was used. The aim was to perform once again the validation of the computational method and to obtain better understanding of the mechanisms of tone noise generation in multistage turbomachines. The results of the investigation were compared with the experimental data obtained in the CIAM C-3A acoustic test facility. In general satisfactory correspondence between the numerical calculations and the experiment was observed. This can be treated as a significant argument for the validity of the frequency domain method of the multistage turbomachinery tone noise calculations.
Proceedings Papers
Proc. ASME. GT2020, Volume 4B: Combustion, Fuels, and Emissions, V04BT04A059, September 21–25, 2020
Paper No: GT2020-16213
Abstract
Wall temperature measurements with fiber coupled online phosphor thermometry were, for the first time, successfully performed in a full scale H-class Siemens gas turbine combustor. Online wall temperatures were obtained during high-pressure combustion tests up to 8 bar at the Siemens CEC test facility. Since optical access to the combustion chamber with fibers being able to provide high laser energies is extremely challenging, we developed a custom-built measurement system, consisting of a water-cooled fiber optic probe and a mobile measurement container. A suitable combination of chemical binder and thermographic phosphor was identified for temperatures up to 1800 K on combustor walls coated with a thermal barrier coating (TBC). To our knowledge these are the first measurements reported with fiber coupled online phosphor thermometry in a full scale high-pressure gas turbine combustor. Details of the setup and the measurement procedures will be presented. The measured signals were influenced by strong background emissions, probably from CO 2 * chemiluminescence. Strategies for correcting background-emissions and data evaluation procedures are discussed. The presented measurement technique enables detailed study of combustor wall temperatures and using this information an optimization of the gas turbine cooling design.
Proceedings Papers
Proc. ASME. GT2020, Volume 7C: Heat Transfer, V07CT14A016, September 21–25, 2020
Paper No: GT2020-14975
Abstract
In order to produce efficient engines it is essential for gas tur-bine designers to understand the interaction between the primary and secondary air systems in critical parts of the engine. One of these is the first stage turbine, where the ingress of the hot an-nulus air into the rotor stator cavity could be catastrophic due to the increased heat load on the disc posts and on the rotor blades themselves (through reduced cooling). To ensure that this does not happen, contactless seals (rim seals) are built into the outer radius of the rotating disc. Additionally, a secondary air flow rate must be appropriately set in order to ‘purge’ the hot air that could be ingested into the rim seal cavity. However, this purge airflow could cause deterioration of the turbine performance as it re-joins the main annulus flow at the interface between the rim seal cavity and the main annulus. The deterioration in performance is pri-marily due to the difference in kinematic (flow velocity and mass flow) and thermodynamic (density, enthalpy) properties of the two stream of air. It is therefore essential to understand the optimum seal geometry and purge flow rates required to prevent the ingestion of the hot annulus air while maintaining the required turbine performance. In this paper we present experimental test results from a single stage turbine facility, the Rim Seal (RiSe) rig, at the University of Sussex. The turbine stage incorporates a model rotor-stator cavity system that is representative of the first stage turbine in a gas turbine engine. The facility is capable of generating disc cavity rotational Reynolds numbers of the order of 2.2 × 10 6 and axial Reynolds number of the order of 0.7 × 10 6 , while operating at a pressure ratio of 2.5. The paper will present the salient features of the test facility, the various instrumentation employed, and the operating specifications of the stage. The paper will discuss the effect of varying the purge flow for a fixed operating point of the turbine. Results presented will include typical mission profiles, cavity radial temperature distribution, and the measured cavity sealing effectiveness.
Proceedings Papers
Proc. ASME. GT2020, Volume 7B: Heat Transfer, V07BT12A058, September 21–25, 2020
Paper No: GT2020-15644
Abstract
A new aerothermal test facility was constructed for the purpose of studying film cooling performance in an environment that accurately simulates conjugate heat transfer characteristics that exist in engine operation. This paper details the design of the facility and the plan for conducting steady-state film cooling experiments to improve the understanding of conjugate heat transfer scaling from laboratory to engine conditions. The test facility consists of two separate flow channels (hot gas/coolant) and each gas path has a flow conditioning section, a convergent nozzle and a test section/channel with viewports. Numerical simulations were conducted to predict flow field characteristics supporting the design of the flow loop facility. Preliminary experiments were conducted to characterize the flow field using velocity and temperature profile measurements. In addition, infrared (IR) thermography methods were developed to measure surface temperatures on the hot side of the test plate. The IR measurement methods including calibration of the IR camera is explained in detail. It was concluded that appropriate hot gas path flow conditioning could be achieved using a strainer-like tube, a perforated plate, and a honeycomb-mesh screen system upstream of the test section. Flow field measurements from preliminary experiments showed that the boundary layer profile follows the law of the wall.
Proceedings Papers
Proc. ASME. GT2020, Volume 8: Industrial and Cogeneration; Manufacturing Materials and Metallurgy; Marine; Microturbines, Turbochargers, and Small Turbomachines, V008T20A009, September 21–25, 2020
Paper No: GT2020-14779
Abstract
An engine test facility, capable of being operated with different nozzles, has been developed. The shape of this component has evolved substantially, over the years, from seemingly simple circular geometries to very complex geometries designed to address different requirements. There are numerous test facilities that facilitate an extensive standalone component level testing of these components. However, such testing does not offer an insight into the effect these complex geometries will have on the performance of a gas turbine engine. This facility will serve as a demonstration for graduate and doctoral students, enhancing their understanding of engine performance. The core of the test facility is a single spool turbojet engine from AMT, Netherlands. The engine is instrumented with pressure and temperature measurements at every inter-component location. A conventional intake duct is designed, for measurement of air flow rate. The entire engine is mounted on a 6-axis force measurement device, for measurement of thrust during engine operations. Towards the rear of the engine, a straight duct is attached after the jet pipe, on which any new nozzle can be retrofitted to the engine. To address changes in operating conditions when using different nozzles, the engine is equipped with a suitably designed bleed duct, which is attached onto the jet pipe. The engine, successfully integrated with all the components, has been subjected to multiple tests at different power settings. The first test was done using the baseline engine, operated at various rotor speeds. The next test was done with a custom nozzle having an area ratio of 2, attached to the jet pipe. Owing to the successful design of the components, the engine performance was measured when operated with a smaller area nozzle.
Proceedings Papers
Proc. ASME. GT2020, Volume 7A: Heat Transfer, V07AT11A013, September 21–25, 2020
Paper No: GT2020-16151
Abstract
In aero engines the combustors are subjected to critical thermal conditions in terms of high temperatures and corrosive environment, which could affect the service life of the entire system. As well known, Thermal Barrier Coatings (TBC) and above all cooling systems represents the state-of-the-art in the nowadays protecting methods: the maximization of this beneficial effect is achieved by defining an optimal cooling arrangement and developing suitable manufacturing technologies for these systems. In modern aero-engine combustors, one of the most effective cooling scheme for liners is composed by an effusion perforation coupled with a slot system to start the film cooling. The cooling performances are deeply influenced by the mutual interactions between swirling and cooling flows. In addition, for typical Rich-Quench-Lean (RQL) combustor architectures, the injection of air provided to promoting the local break-down of the flame mixture fraction, deeply interacts with the swirled flow, generating recirculating structures capable of affecting the development of film cooling and making the design of cooling systems very challenging. A new test facility for testing effusion test plates for RQL combustors applications has been developed with the final aim of comparing different cooling strategies and at the same time to collect data for numerical model validation. The experimental set-up consists of a non-reactive planar sector rigs with 5 engine-scale swirlers fed with air up to 250 °C and 3 bar. The rig was equipped with outer/inner dilution ports, and a simple inner liner cooling scheme composed of effusion and a slot system: all these features, fed with air at ambient temperature, can be independently controlled in terms of mass flow. Using dedicated optical accesses, InfraRed (IR) camera tests were performed to retrieve overall effectiveness data imposing a temperature difference between swirling and cooling flows. To better understand those results, Pressure Sensitive Paint (PSP) technique was used to obtain reliable film effectiveness data decoupling the contribution of slot and effusion flows. The thermal characterization was supported by Particle Image Velocimetry (PIV) investigations on the median plane. Tests were performed at different pressure drops across swirler and varying the mass flows of slot and inner/outer liners. The analysis of the data highlighted the influences of the swirling flow on the overall thermal performance and the behaviour of the film cooling system.
Proceedings Papers
Proc. ASME. GT2020, Volume 7C: Heat Transfer, V07CT13A009, September 21–25, 2020
Paper No: GT2020-14528
Abstract
A high-speed infrared camera is used to measure the temperature of blade tips in a cooled high-pressure turbine operating at corrected engine conditions in The Ohio State University short duration Turbine Test Facility. These experiments create a challenging problem for infrared imaging since the rotor turns at over 13,000 rpm with tip speeds on the order of 300 m/s, and the surface temperature of the airfoils is on the order of 350 K. This means that the camera needs to capture a low intensity signal in a very short time period. This paper will review the design and operation of a measurement procedure to accomplish this difficult task along with the post-processing steps necessary to extract useful data. Raw infrared images are processed by deblurring the images using a non-blind Wiener filter and mapping the two-dimensional data onto the three-dimensional blade. This paper also describes experiments covering a range of cooling flow rates and main flow temperatures. In addition, several tests with no main flow and only cooling flow were performed at lower speeds to reduce motion blur and enable the separation of internal and external heat transfer information. Results show that the infrared data is consistent and can provide quantitative comparisons of cooling performance even at the high rotation speed. This paper presents the lessons learned for high-speed infrared measurement along with representative data to illustrate the repeatability and capability of the measurement scheme as well as suggested improvements to guide further development.
Proceedings Papers
Proc. ASME. GT2020, Volume 11: Structures and Dynamics: Structural Mechanics, Vibration, and Damping; Supercritical CO2, V011T31A002, September 21–25, 2020
Paper No: GT2020-14334
Abstract
A team led by Gas Technology Institute (GTI ® ), Southwest Research Institute ® (SwRI ® ) and General Electric Global Research (GE-GR), along with the University of Wisconsin and Natural Resources Canada (NRCan), is actively executing a project called “STEP” [ S upercritical T ransformational E lectric P ower project], to design, construct, commission, and operate an integrated and reconfigurable 10 MWe sCO 2 [supercritical CO 2 ] Pilot Plant Test Facility. The $122* million project is funded $84 million by the US DOE’s National Energy Technology Laboratory (NETL Award Number DE-FE0028979) and $38* million by the team members, component suppliers and others interested in sCO 2 technology. The facility is currently under construction and is located at SwRI’s San Antonio, Texas, USA campus. This project is a significant step toward sCO 2 cycle based power generation commercialization and is informing the performance, operability, and scale-up to commercial plants. Significant progress has been made. The design phase is complete (Phase 1) and included procurements of long-lead time deliver components. Now well into Phase 2, most major equipment is in fabrication and several completed and delivered. These efforts have already provided valuable project learnings for technology commercialization. A ground-breaking was held in October of 2018 and now civil work and the construction of a dedicated 25,000 ft2 building has progressed and is largely completed at the San Antonio, TX, USA project site. Supercritical CO 2 (sCO 2 ) power cycles are Brayton cycles that utilize supercritical CO 2 working fluid to convert heat to power. They offer the potential for higher system efficiencies than other energy conversion technologies such as steam Rankine or Organic Rankine cycles this especially when operating at elevated temperatures. sCO 2 power cycles are being considered for a wide range of applications including fossil-fired systems, waste heat recovery, concentrated solar power, and nuclear power generation. By the end of this 6-year STEP pilot demo project, the operability of the sCO 2 power cycle will be demonstrated and documented starting with facility commissioning as a simple closed recuperated cycle configuration initially operating at a 500°C (932°F) turbine inlet temperature and progressing to a recompression closed Brayton cycle technology (RCBC) configuration operating at 715°C (1319 °F).
Proceedings Papers
Proc. ASME. GT2020, Volume 11: Structures and Dynamics: Structural Mechanics, Vibration, and Damping; Supercritical CO2, V011T31A018, September 21–25, 2020
Paper No: GT2020-15945
Abstract
sCO 2 power cycles offer improved cycle efficiencies compared with traditional steam Rankine cycles. However, the turbomachinery required to support such a cycle does not exist at a commercial scale and requires development. This paper describes a new 10 MWe scale sCO 2 turbine was developed and demonstrated in an sCO 2 closed-loop recompression Brayton cycle. Since this turbine was developed for Concentrating Solar Power (CSP) applications, a target inlet temperature of over 700°C was chosen using funding from the US DOE SunShot initiative and industry partners. However, it can be applied to traditional heat sources such as natural gas, coal, and nuclear power. Traditional Rankine steam cycle thermal efficiencies are typically in the 35–40% range, but can be as high as 45% for advanced ultra-supercritical steam cycles. The sCO 2 cycle can approach 50% thermal efficiency using externally fired heat sources. Furthermore, this cycle is also well suited for bottoming cycle waste heat recovery applications, which typically operate at lower temperatures. The high-power density and lower thermal mass of the sCO 2 cycle results in compact, high-efficiency power blocks that can respond quickly to transient environmental changes and transient operation, a particular advantage for solar, waste heat, and ship-board applications. The power density of the turbine is significantly greater than traditional steam turbines and is comparable to liquid rocket engine turbo pumps. This paper describes the design and construction of the turbine and provides additional testing of the 10 MWe turbine in a 1 MWe test facility including a description of rotordynamics, thermal management, rotor aero and mechanical design, shaft-end and casing seals, bearings, and couplings. Test data for the turbine is included, as it achieves its operational goal of 715°C, 250 bara, and 27,000 rpm.
Proceedings Papers
Proc. ASME. GT2020, Volume 2A: Turbomachinery, V02AT32A068, September 21–25, 2020
Paper No: GT2020-16104
Abstract
Advancements in core compressor technologies are necessary for next generation, high Overall Pressure Ratio (OPR) turbofan engines. High pressure compressors (HPCs) for future engines are being designed with exit corrected mass flow rates less than 2.25 kg/s (5 lbm/s). In order to accurately measure the performance of these advanced designs, high accuracy measurements are needed in test facilities. The W7 High Speed Multistage Axial Compressor Facility at NASA Glenn Research Center has been used to acquire data for advanced compressor designs. This facility utilizes an advanced differential pressure flow meter called a V-Cone. The facility has historically tested components with physical mass flow rates in the range of 27 to 45 kg/s (60 to 100 lbm/s). As such, when the V-Cone was calibrated prior to installation, the calibrations focused on higher mass flow rates, and uncertainties in that regime range from 0.5% to 0.85%. However, for low mass flow rates under 9 kg/s (20 lbm/s), expected in tests of advanced high OPR HPCs rear stages, the uncertainties of the V-Cone exceed 2.5%. To address this, using a method similar to that utilized by the National Institute of Standards and Technology, an array of Critical Flow Venturi Nozzles (CFVs) was installed in the W7 test section and used to calibrate the V-Cone in 0.5 kg/s (1 lbm/s) increments up to 10.5 kg/s (23 lbm/s). This effort details the measurements and uncertainties associated with this calibration which resulted in a final uncertainty of the V-Cone measurements under 1%.
Proceedings Papers
Proc. ASME. GT2020, Volume 9: Oil and Gas Applications; Organic Rankine Cycle Power Systems; Steam Turbine, V009T23A028, September 21–25, 2020
Paper No: GT2020-16307
Abstract
Conventional power plants are obliged to compensate for the fluctuations in power generation, due to the rising amount of renewable energies, to ensure grid stability. Consequently, steam turbines are more frequently facing load variation and startup/shut-down cycles leading to an increase of thermal stress induced by phase change phenomena. The review of existing test facilities providing measurement data of heat transfer coefficients influenced by multiphase phenomena, such as surface wettability and dry-out, revealed the necessity for a new measurement application. This paper presents the design of the Experimental Multi-phase Measurement Application “EMMA” to generate the required conditions in combination with an academic turbine housing geometry. The performed investigations are focused on the local distribution of heat transfer coefficients (HTC) and the surface wettability affected by phase change phenomena. Two main film formation mechanisms can be observed, depending on the thermal gradient between the fluid and the wall. These are a) saturated/superheated steam in contact with a sub-cooled wall leading to film-wise/drop-wise condensation and b) primary condensed wet steam droplets depositing on a superheated wall, leading to evaporation. Both, the liquid film and the local heat transfer are measured simultaneously. An overview of applicable thickness measurement methods for transparent liquid films is given and the applied optical measurement system is further described. Moreover the HTC measurement methods are presented considering the occurring case of phase change.
Proceedings Papers
Proc. ASME. GT2020, Volume 10A: Structures and Dynamics, V10AT25A012, September 21–25, 2020
Paper No: GT2020-14464
Abstract
Dry gas lubricated non-contacting mechanical seals (DGS), most commonly found in centrifugal compressors, prevent the process gas flow into the atmosphere. Especially when high speed is combined with high pressure, DGS is the preferred choice over other sealing alternatives. Even though the non-contacting seal is proved reliable; the ultra-thin gas film can still lead to a host of potential problems due to possible contact. In order to investigate the flow field in the sealing gap and to facilitate the numerical prediction of the seal performance, a dedicated test facility is developed to carry out the measurement of key parameters in the gas film. Gas in the sealing film varies according to the seal inlet pressure, and the thickness of gas film depends on this fluctuated pressure. In this paper, the test facility, measurement methods and the first results of static pressure measurements in the sealing gap of the DGS obtained in the described test facility are presented. An industry DGS with three-dimensional grooves on the surface of the rotating ring, where experimental investigations take place, is used. The static pressure in the gas film is measured, up to 20 bar and 8,100 rpm, by several high frequency ultraminiature pressure transducers embedded into the stationary ring. The experimental results are discussed and compared with the numerical model programmed in MATLAB [1], the characteristic and magnitude of which have a good agreement with the numerical simulations. It suggests the feasibility of measuring pressure profiles of the standard industry DGS under pressurized dynamic operating conditions without altering the key components of the seal and thereby affecting the seal performance.
Proceedings Papers
Proc. ASME. GT2020, Volume 10A: Structures and Dynamics, V10AT25A027, September 21–25, 2020
Paper No: GT2020-15520
Abstract
The design, construction, operational capabilities, and proof of concept results are presented for a test rig used to evaluate gas-lubricated thrust bearings. The following work is motivated by a desire to utilize the working fluid of high-performance turbomachinery, such as gas turbines, for bearing lubricant. Auxiliary equipment required to cool, pump, and clean oil for a typical thrust bearing is eliminated by taking advantage of the turbomachinery’s working fluid as bearing lubricant. The benefit of removing such auxiliary equipment is obvious when considering cost and weight of turbomachines, yet the working fluid of gas turbines typically has very low viscosity compared to oil which introduces load capacity and stability challenges. It is therefore necessary to build a facility capable of testing gas-lubricated thrust bearings to advance the technology. The test rig design in this work allows for 7 to 15 inch (180–380 mm) diameter thrust bearings, static loads up to 30,000 lbf (135 kN), and speeds up to 20 krpm. The test facility also provides up to 500 psig (3.45 MPa) static air pressure to enable testing of hydrostatic and hybrid (hydrodynamic combined with hydrostatic) bearings. This paper describes the test rig operating principle, details experimental procedures to obtain measurements, and provides test results necessary to prove the test rig concept by means of a hybrid gas bearing.
Proceedings Papers
Proc. ASME. GT2020, Volume 10A: Structures and Dynamics, V10AT24A009, September 21–25, 2020
Paper No: GT2020-14915
Abstract
A composite fan stage representative of a modern UHBR architecture has been investigated experimentally on a novel test facility at Ecole Centrale de Lyon. These measurements show indications for strong overloading of the tip region resulting in extensive blockage of the blade passage. The performance of the fan is analyzed with extensive instrumentation including radial profiles upstream and downstream of the rotor. Unsteady pressure measurements help to interpret the flow structure in the tip region. The results are presented across a range of operating points on the design speedline. At the stability limit, the machine suffers from Non-Synchronous Vibrations which result from small scale aerodynamic disturbances propagating between the leading edges. A detailed analysis on the occurring waveforms is presented for two operating speeds. In order to further analyze the observed phenomena, a numerical study has been conducted using the RANS solver elsA. The results of steady calculations are discussed in comparison with the detailed experiments. Unsteady simulations near the stability limit accurately predict the aerodynamic disturbances observed during NSV. The obtained results are unusual for typical state-of-the-art transonic fans, as they show the same behavior as high-pressure compressor front stages, dominated by blockage caused by tip leakage flow. Even though flutter is not observed, the observed Non-Synchronous Vibration mechanism is a critical aeroelastic phenomenon which is of great interest for future designs of low speed fans.
Proceedings Papers
Lorenzo Pinelli, Federico Vanti, Lorenzo Peruzzi, Andrea Arnone, Andrea Bessone, Claudio Bettini, Roberto Guida, Michela Marré Brunenghi, Vaclav Slama
Proc. ASME. GT2020, Volume 10A: Structures and Dynamics, V10AT24A017, September 21–25, 2020
Paper No: GT2020-15409
Abstract
This paper is part of a two-part publication that aims to experimentally and numerically evaluate the aerodynamic and mechanical damping of a last stage ST blade at low load operation. A three-stage downscaled steam turbine with a snubbered last stage moving blade LSMB has been tested in the T10MW test facility of Doosan Skoda Power R&D Department in the context of the FLEXTURBINE European project (Flexible Fossil Power Plants for the Future Energy Market through new and advanced Turbine Technologies). Aerodynamic and flutter simulations of different low load conditions have been performed. The acquired data are used to validate the unsteady CFD approach for the prediction of the aerodynamic damping in terms of logarithmic decrement. Numerical results have been achieved through an upgraded version of the URANS CFD solver, selecting appropriate and robust numerical setups for the simulation of very low load conditions, such as increased condenser pressure at the exhaust hood outlet. The numerical methods for blade aerodamping estimation are based on the computation of the unsteady pressure response caused by the row vibration. They are usually classified in time-linearized, harmonic balance and non-linear approaches both in frequency and time domain. The validation of all these methods historically started in the field of aeronautical low-pressure turbines and has been gradually extended to compressor blades and steam turbine rows. For the analysis of a steam turbine last rotor blade operating at strong part load conditions, non-linear methods are recommended as these approaches are able to deal with strong nonlinear phenomena such as shock waves and massive flow separations inside the domain. Experimental data have been used to separate the contributions of mechanical and aerodynamic damping, extrapolating to zero mass flow the total measured damping. Finally, the comparisons between the aerodynamic damping coming from measurements and CFD results have been reported in order to highlight the capability to properly predict the last stage blade flutter stability at low load conditions. Such comparisons confirms the flutter free design of the new snubbered LSMB blade.
Proceedings Papers
Proc. ASME. GT2020, Volume 9: Oil and Gas Applications; Organic Rankine Cycle Power Systems; Steam Turbine, V009T21A019, September 21–25, 2020
Paper No: GT2020-16332
Abstract
Hydrogen is one of the leading options for storing energy from renewables and surplus electricity. Hydrogen is also a major constituent in various streams in the chemical industry and cannot always be used for better purposes than flaring or heat and power generation. Siemens has identified the 24MWe SGT-600 3 rd generation DLE gas turbine as a candidate for having a high hydrogen capability. The burners for using hydrogen in the SGT-600 have been developed for and by Additive Manufacturing technology. The advantages of this technology have been integrated into the presented design and therefore allowing: • Rapid prototyping with possibilities for fast turnaround of tests and screening of various concepts • Manufacturing of complex geometries with smart gas passages, very innovative cooling and mixing concepts • Small series and minimum waste with reduced cost • Good repeatability and stability of product quality Burner development was carried out according to “the standard method within the industry”, meaning CFD-analysis, atmospheric single burner combustion testing followed by pressurized single burner combustion testing and finally a full-scale machine test at the SIEMENS Industrial Turbomachinery AB (SIT) test rig facility in Sweden. The rig is used for full scale testing of gas turbines in the power output range from 15MW to 62MW. It allows testing not only with standard natural gas but also gas mixtures with e.g. hydrogen or nitrogen can be run. The test facility has liquid fuel capability. During the burner development process, a project including two SGT-600 running on up to 60 volume % hydrogen was awarded to Siemens. This meant that a very definite target for the development was set and the results of these efforts are presented in this paper. An adapted 3 rd gen. DLE burner design proved to be capable of using 100% hydrogen at SGT-600 full load conditions at the single burner high pressure tests giving only 35 ppm NO x @15%O 2 . This was a major step in the development of a hydrogen burner for the SGT-600. The following full engine test with the same burner type showed the possibility to run with 60 vol-% H 2 at 0–100% load while keeping stable combustion and achieving emissions below 25 ppm NO x @15%O 2 in the standard operating range of the SGT-600. At lower loads higher hydrogen contents were tested (95 vol-%) but the flow capacity of the fuel system limited the full exploration of hydrogen capability of the SGT-600 3 rd gen. DLE gas turbine. The 3 rd gen. DLE burner is also used in the 33MWe SGT-700 and the 62MWe SGT-800, which will also benefit from the development of the increased SGT-600 hydrogen capability. The results open the possibility of using H 2 rich gas more widely in all gas turbine configurations using 3 rd gen. DLE burner.
Proceedings Papers
Proc. ASME. GT2020, Volume 10A: Structures and Dynamics, V10AT24A019, September 21–25, 2020
Paper No: GT2020-15450
Abstract
This paper is the first of a two-part publication that aims to experimentally evaluate, simulate and compare the aerodynamic and mechanical damping for a last stage steam turbine rotor blade at part load operation. Resulting strong off-design partial load regimes expose the last stage moving blade (LSMB) to the possible onset of aero-elastic instabilities, such as stalled and un-stalled flutter. This interaction can lead to asynchronous blade vibrations and then the risk of blade failures for high cycle fatigue. In this framework, it is necessary to develop and validate new tools for extending operating ranges, controlling non-synchronous phenomenon and supporting the design of new flutter resistant LSMB. To this end, a 3-stage downscaled steam turbine with a snubbered LSMB was designed by Ansaldo Energia and tested in the T10MW test facility of Doosan Skoda Power R&D Department within the FlexTurbine European project. The turbine was operated in a wet steam environment at very low volume flow conditions simulating different part load regimes. The steady flow field throughout the LSMB was characterized and the occurrence of flutter was investigated by inducing the blade resonance through an AC magnet excitation and measuring the overall damping. The results presented in this paper indicate that the blade always operates over the flutter stability margin validating this new blade design. In the second part of this work, the mechanical and aerodynamic contribution to the damping will be separated in order to validate the aerodynamic damping prediction of an upgraded CFD tool, already adopted in the design phase of the blade at design point.